1. Field of the Invention
The following invention relates generally to high temperature gas valves and more particularly to a technique for reducing the effects of thermal shock upon a high temperature gas valve.
2. Related Art
Vehicles that propel through outer space must withstand extreme conditions, unlike any encountered on earth. For example, a space vehicle must withstand temperatures near absolute zero (zero degrees Kelvin), while burning fuels at extremely high temperatures (in the thousands of degrees) for attitude control, which is the control of the vehicle position and velocity in outer space.
The way a space vehicle moves in a solid propellant-based system is as follows. A propellant is burned (which has mass) generating gas (that has mass with a velocity). The velocity direction of ejection is controlled through a series of manifolds, leading to control valves and thrust systems. The valves control the amount of propellant and the direction of its travel, thus controlling the attitude of the vehicle. Very precise control must be maintained over how much impulse is delivered in any given direction. These decisions are made by an onboard flight control computer that decides what corrections are necessary, and the computer then communicates those signals electrically to the control valves.
One great problem is that the control valves, which are advantageously constructed of thermal shock sensitive materials, are exposed to the extreme temperature differentials between the hot inlet gas and the cold of outer space. The valves are constructed of solid materials which are required to have significant physical strength at the temperatures at which they operate.
The trend in recent years has been to adopt higher and higher flame temperatures (to heat the propellant) in order to get a propulsive efficiency gain. Higher flame temperatures allow the packaging of more energy in a given pound of propellent and, therefore, greater thrust per a given space (or mass metric) in a satellite or missile system space vehicle.
There is a certain amount of space or mass allocated to an attitude control system (or a main thrust propulsive system). To the extent that the propulsion designer (the designer of the gas generator or the rocket motor) can raise the flame temperature of the combusting propellent system, the more energy can be packed into a given volume or mass. This technology trend has been ongoing since the beginning of the propulsion history.
What limits how high a temperature such systems can achieve is generally the performance limits of the physical devices that are connected to the rocket motor or the gas generator, which are for example, the valve and the manifold system in the attitude control system and the rocket nozzle, associated with the main rocket motor propulsion type system.
As noted, thermal shock is created by the differential between the hot inlet propellant gases and the cold of outer space. The temperature difference between the core of the material (of the valve for example) and the surface of the material results in very high stresses, because the surface rapidly expands or rapidly contracts depending on the direction of the temperature change. The underlying core material has little time to react because the temperature pulse has not reached it. The change in expansion or contraction creates a stress field which is destructive to the material of the device.
Thermal shock is most destructive for brittle materials that cannot stretch very far to accommodate the temperature change without failing, or rupturing, or cracking. This thermal shock sensitive behavior occurs most often in physical devices where brittle materials (such as graphite and ceramics) are employed to withstand high temperature changes because these materials are most resistant. Such brittle materials are employed for control valves, which are subjected to very high rates of heating.
The problem of thermal shock has been attacked in a number of ways. One method is to select different materials that are more ductile, which stretch rather than crack when these thermally induced strains are forced on them by the temperature differences. However, this solution is frequently not workable for rocket propulsion space vehicles, where temperatures range between 3,000-5,000xc2x0 F., because the propellant temperatures are so high that few materials have usable strength at these temperatures.
At these temperatures, carbon materials (e.g., diamond, graphite) and certain ceramics compositions (e.g., nitrides and carbides) and a few varieties of metals (e.g., tungsten, tantalum, molybdenum) are usable. Many of these materials are brittle, making them quite thermal shock sensitive.
A second solution is to decrease the amount temperature of the inlet propellant gas. However, as noted, extreme temperatures differentials are needed for modem space vehicles, to promote efficiencies of the propellant system.
A third solution is to decrease the time rate of the change of temperature, which is different than just the actual temperature, itself. Taking a material to a very high or very low temperature quickly exaggerates the difficulty of the thermal shock challenge to the material. There are two competing rates that take place in the physics of the thermal shock. One is the rate at which heat is either donated to or liberated from (extracted from) the surface of the material.
The second rate is the rate at which the heat is internally redistributed within the material. To the extent that temperature is delivered extremely rapidly or heat is delivered extremely rapidly to the surface of the part and not removed (or dissipated) to the interior of the part by conduction at an equally rapid rate, the heat accumulates spatially at the surface layers of the part and, therefore, creates a much stronger gradient (or rate of change) on the material part. This produces very high changes in thermally-induced strains, and results in very high stress fields (or very high stress differences) from one layer in the part to the next layer.
If that same temperature rise is delivered over a longer period of time, the two competing rates will end up having a different balance. The first conduction rate that pulls the heat away from the surface and redistributes it will offset the second rate, thus tending to equalize the temperature of the part throughout the thickness of the part. This gives the material of the part more time to cope with the influx of heat at the surface and, therefore, does not accumulate or build up a large temperature gradient throughout the thickness of the part. The reason is because the internal conduction mechanism has more time to pull the heat away and prevents it from literally accumulating at the surface.
What is required is a satisfactory method to reduce the rate of temperature change to reduce the level of thermal shock stresses, in particular for brittle material parts used to cope with sudden exposure to extreme temperatures. This method, and system for employing the method, must reduce the rate at which heat is donated to the surface of the material, to more closely equal the rate at which the heat is internally redistributed within the material.
The present invention is directed to a system (and accompanying method) for reducing thermal shock in a physical device, including: a gas inlet portion; and an ablator material positioned in the gas inlet portion to reduce a temperature differential at a surface between the gas inlet portion and a gaseous material flowing through the gas inlet portion. The gas inlet portion can include: a first material; and a second material having a boundary with the first material, where the ablator material is positioned at the boundary. In one embodiment, the first material is not sensitive to thermal shock, and the second material is sensitive to thermal shock. The ablator material can be housed in a recess at the boundary.
The first material can be a manifold. The second material can be a portion of a control valve for controlling the direction and amount of a hot gas emitted therefrom. The manifold can include a bypass slot connecting a main inlet gas chamber of the manifold with a surface of the ablator material.
A portion of the manifold between the bypass slot and the main inlet gas chamber can be diverted to block a portion of the volume of the main inlet gas chamber. A portion of the ablator material can protrude into the main inlet gas chamber. A portion of the manifold can protrude into the main inlet chamber to form an eddy gas current in contact with the ablator material and behind the manifold. A portion of the ablator material can also protrude into the main inlet gas chamber.
The system can further include a consumable ring protruding into the main inlet gas chamber and positioned in contact with the ablator material, where the consumable ring is comprised of a material which consumes less rapidly than the ablator material. The ablator material can protrude into the main inlet gas chamber.
The ablator material can be comprised of a material having little residue upon heating and if possible, one which changes directly from a solid to a gaseous form without first changing to a liquid intermediate form. Specifically, the ablator material can be comprised of any one of: a paraffin material; a polyethylene material; a thermoplastic material; an octadecane material; and a phenolic material. The ablator material can also be comprised of a first material and any one of: a graphite powders filler; a talcum powder filler; and a phase change salt filler.